Conventional rocket motors use a propellant grain which is normally cast in place. The burning surface of the grain is the exposed surface of the cast grain and is limited by the grain geometry. Structural integrity of the grain is affected directly by rocket acceleration which tends to crack the grain, deflect it and separate the grain from the case wall at the bond line. The motor case may have a rubber type of liner bonded to its interior with the propellant grain, in turn, bonded to the liner.
Conventional rockets are stress limited to maximum accelerations of about 500 g and minimum burn times of about 0.5 seconds. For example, large ABM rocket motors such as HiBex have been built with a propellant burning rate of 4 inches per second at 2,000 psi. Small scale experiments have shown that a burning rate of 12 inches per second is possible, however, the burning rate affects volumetric loading efficiency and does not affect grain stress. HiBex test vehicles have reached a velocity of 5,000 fps in 0.8 seconds at a range of about 1,100 feet and attained a maximum velocity of about 9,000 fps in 1.125 seconds at a range of 3,000 feet with a 750 g maximum acceleration at burnout.
In these "conventional" rocket motors, a single hollow cast propellant grain, approximately the consistency of rubber, is bonded to the motor case or to a synthetic rubber liner which in turn is bonded to the motor case and only the inner surface of the grain burns. The propellant grain is put under high stress due to its inertia during rocket acceleration and due to internal pressure expansion of the motor case which tends to crack the grain and separate it from the case wall. In the ABM rocket motors, the thick grain must withstand axial acceleration in the order of 750 g's. The overall propellant stresses and strains govern acceptable motor case and propellant thickness which result in motor mass fractions of approximately 0.8. Space rockets may have a mass fraction of 0.9, mass fraction being generally defined as the weight of propellant divided by the launch weight of the rocket.
There are several methods of increasing rocket motor performance with the current state of the art technology. One approach is to design a motor in which the nozzle is part of the propellant grain resulting in higher mass fractions due to the reduction of motor hardware weight. This solution does nothing for the propellant/bond stress and strain problem and does not decrease motor burn time.
Another aspect of rocket motor design and the problems with prior motor designs is related to the use of a rocket motor as a device to deliver accurately to a target a kinetic energy penetrator or other type of payload at a relatively high and preferably hypersonic velocity. A typical such device is an anti-tank impulse rocket or one used to deliver a high energy penetrator to a reinforced structure such as a building or bunker or to deliver a battering ram type of device to building for rapid access thereto (the battering ram requires velocities only in the 1,000 ft/sec regime but burnout distances are restricted to 5 feet or less so that the battering ram can be placed in restricted areas).
For example, current antitank rockets are generally subsonic rockets with burning times of greater than one second. These are high explosive rounds with maximum ranges on the order of one or two miles at maximum velocity. The velocity of the explosive warhead antitank rocket does not add to the penetration effectivity.
Where the mission requirement requires close in use, then the rocket system should have a burn time in the order of milliseconds to achieve a kill velocity, e.g., 6,000 fps, at point blank range. An explosive warhead would probably injure or kill the launching personnel at target ranges less than 50 feet. Thus for example, calculations (assuming a constant acceleration) indicate that the burn time is between 10 to 100 milliseconds to reach a target with a velocity of 6,000 fps where the target ranges are from 20 to 300 feet. Conventional and known rocket motors have not been able to achieve these velocities in a short burn time.
Such relatively short burn times produce a weapon that can get to a distant target in a very short time so that allowance for target motion is minimized in ranges in the order of 1,000 meters. It is known, for example, that the flight error is related to burn time. Thus, a shorter burn time to achieve a relatively high velocity tends to reduce flight error. Assuming a rocket travel of 1,000 meters in 0.55 seconds, a tank travelling normal to the line of sight of the rocket launch position at 20 mph (32 km/h) would move only 18.8 feet after firing and before being hit by a 6,000 fps munition. Because such an impulse rocket accelerates to full velocity in less than 50 feet, minimal or no allowance need be made for wind drift and weathercocking angle of attack.
It is also apparent that not all impulse rockets need the high velocity normally required for tank armor penetration. The velocity of the impulse rocket of this invention may be between 500 and 10,000 fps in 100 milliseconds or less. There are instances in which it is desired to use a rocket device, close in, in order to deliver a debilitating gas to the interior of the structure or to form an opening in a structure without creating a fire and without injuring those in the structure, i.e., a hostage assault and recovery operation. In this instance, the velocity of the rocket should be less than 6,000 fps so as to prevent the payload from passing all the way through the structure-from one side to the other. Typically a velocity of between 1,000 to 2,000 fps and as high as 5,000 fps is adequate for these purposes, depending on the needs of the mission. Particularly advantageous is the relatively short burn time and the achievement of a relatively high velocity in a short travel interval to provide a high energy impact. The energy of impact is related to the kinetic energy, i.e., 1/2(mass.times.velocity.sup.2). In this instance, the rocket motor is no longer burning when the projectile impacts the wall and thus, even if the rocket and projectile enter the structure after taking out a wall or forming an opening therein, there is no possibility of a rocket induced fire within the structure.
It is thus apparent that a need exists for an efficient rocket motor with burn times of less than about 0.010 seconds.
It is equally apparent that a need exists for a high energy impulse rocket, i.e., a rocket system that develops a large kinetic energy over a relatively short period of time.
It is also clear that a need exists for an improved rocket motor in which grain cracking, grain stress, and grain strain no longer limit motor acceleration while at the same time providing a mass fraction as high as possible to obtain maximum velocity with minimal weight and size.
There are distinct advantages to be achieved and significant cost reductions if, for example, cast or preformed propellants and bonding of the same to the rocket motor case can be avoided. These and other practical advantages to be discussed, take on added significance if the rocket motor can achieve short burn times in the range of up to about 100 milliseconds and which can accelerate at g levels ranging from 5,000 g to 20,000 g.
It is also desirable to provide a high velocity impulse rocket which can be fired from a launch tube or other device such as a gun barrel, and which may be mounted or hand held.
Moreover, there are unique advantages to a high velocity, low cost impulse rocket system for use as a sounding rocket, and a low cost orbital insertion system for small payloads. Such systems, in accordance with this invention, may also be used for research requiring high speeds in the atmosphere and on rocket sleds. In addition, the rocket booster of this invention may be used on naval vessels to fire anti-aircraft and anti-missile projectiles without generating a long visible visual or radar/IR trail leading to the ship or launch installation and thereby enhance defense against airborne attack vehicles.